Individual inlet guide vane control for tip turbine engine

A tip turbine engine according to the present invention includes a plurality of independently variable inlet guide vanes for the fan and/or for the compressor. An actuator is operatively coupled to each of the flaps, such that each actuator can selectively vary the flap of its associated inlet guide vane. In one embodiment, the inlet guide vanes each include a pivotably mounted flap that is variable independently of the flaps of at least some of the other inlet guide vanes. In another embodiment, the inlet guide vanes each include at least one fluid outlet or nozzle directing pressurized air, as controlled by the associated actuator, to control inlet distortion.

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Description
CROSS REFERENCE TO RELATED APPLICATIONS

This is a divisional application of U.S. patent application Ser. No. 11/719,868 dated May 22, 2007, now U.S. Pat. No. 8,641,397.

BACKGROUND OF THE INVENTION

The present invention relates to turbine engines, and more particularly to individually controlled inlet guide vanes for a tip turbine engine.

An aircraft gas turbine engine of the conventional turbofan type generally includes a forward bypass fan, a low pressure compressor, a middle core engine, and an aft low pressure turbine, all located along a common longitudinal axis. A high pressure compressor and a high pressure turbine of the core engine are interconnected by a high spool shaft. The high pressure compressor is rotatably driven to compress air entering the core engine to a relatively high pressure. This high pressure air is then mixed with fuel in a combustor, where it is ignited to form a high energy gas stream. The gas stream flows axially aft to rotatably drive the high pressure turbine, which rotatably drives the high pressure compressor via the high spool shaft. The gas stream leaving the high pressure turbine is expanded through the low pressure turbine, which rotatably drives the bypass fan and low pressure compressor via a low spool shaft.

Although highly efficient, conventional turbofan engines operate in an axial flow relationship. The axial flow relationship results in a relatively complicated elongated engine structure of considerable length relative to the engine diameter. This elongated shape may complicate or prevent packaging of the engine into particular applications.

A recent development in gas turbine engines is the tip turbine engine. Tip turbine engines include hollow fan blades that receive core airflow therethrough such that the hollow fan blades operate as a high pressure centrifugal compressor. Compressed core airflow from the hollow fan blades is mixed with fuel in an annular combustor, where it is ignited to form a high energy gas stream which drives the turbine that is integrated onto the tips of the hollow bypass fan blades for rotation therewith as generally disclosed in U.S. Patent Application Publication Nos.: 20030192303; 20030192304; and 20040025490. The tip turbine engine provides a thrust-to-weight ratio equivalent to or greater than conventional turbofan engines of the same class, but within a package of significantly shorter length.

In some applications, there may be a significant component of the airflow that is normal to the inlet to the turbine engine. This normal component may cause distortion of the airflow and cause stability problems. This would be particularly true where the turbine engine is mounted vertically in the aircraft and another engine provides forward thrust. The aircraft would often be moving in a direction normal to the inlet to the vertically-oriented turbine engine. It should be noted that even engines that are not completely vertical may also have a significant component of the airflow that is normal to the turbine engine axis.

SUMMARY OF THE INVENTION

A tip turbine engine according to the present invention includes a plurality of independently variable inlet guide vanes for the fan and/or for the compressor. An actuator is operatively coupled to each of the flaps, such that each actuator can selectively vary the flap of its associated inlet guide vane. In one embodiment, the inlet guide vanes each include a pivotably mounted flap that is variable independently of the flaps of at least some of the other inlet guide vanes. In another embodiment, the inlet guide vanes each include at least one fluid outlet or nozzle directing pressurized air, as controlled by the associated actuator, to control inlet distortion.

With independent control of the variable inlet guide vanes, distortion at the inlet to the bypass fan and/or the inlet to the compressor is reduced, thereby improving the stability of the turbine engine. The independently variable inlet guide vanes can be used in tip turbine engines and other turbine engines. Although potentially useful for horizontal installations as well, this feature is particularly suited for non-horizontal installations, especially vertical installations, where there is a substantial airflow component normal to the inlet to the turbine engine.

BRIEF DESCRIPTION OF THE DRAWINGS

Other advantages of the present invention can be understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein:

FIG. 1 is a longitudinal sectional view along an engine centerline of a tip turbine according to the present invention.

FIG. 2 schematically illustrates three of the fan inlet guide vanes and three of the compressor inlet guide vanes of the tip turbine engine of FIG. 1.

FIG. 3 schematically illustrates the tip turbine engine of FIG. 1 installed vertically in an aircraft.

FIG. 4 illustrates an alternative variable fan inlet guide vane for the turbine engine of FIGS. 1-3.

FIG. 5 illustrates an alternative variable compressor inlet guide vane for the turbine engine of FIGS. 1-3.

 DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

FIG. 1 is a partial sectional view of a tip turbine engine (TTE) type gas turbine engine 10 taken along an engine centerline A. Although the turbine engine 10 is shown horizontally, the turbine engine 10 could be mounted at any orientation, and as explained above, vertical orientations would experience particular benefits from the present invention. The turbine engine 10 includes an outer housing 12, a rotationally fixed static outer support structure 14 and a rotationally fixed static inner support structure 16. A plurality of fan inlet guide vanes 18 are mounted between the static outer support structure 14 and the static inner support structure 16. Each fan inlet guide vane 18 includes a variable flap 18A.

A nosecone 20 may be located along the engine centerline A to improve airflow into an axial compressor 22, which is mounted about the engine centerline A behind the nosecone 20. The nosecone 20 might not be used in vertical installations.

A fan-turbine rotor assembly 24 is mounted for rotation about the engine centerline A aft of the axial compressor 22. The fan-turbine rotor assembly 24 includes a plurality of hollow fan blades 28 to provide internal, centrifugal compression of the compressed airflow from the axial compressor 22 for distribution to an annular combustor 30 located within the rotationally fixed static outer support structure 14.

A turbine 32 includes a plurality of tip turbine blades 34 (two stages shown) which rotatably drive the hollow fan blades 28 relative a plurality of tip turbine stators 36 which extend radially inwardly from the rotationally fixed static outer support structure 14. The annular combustor 30 is disposed axially forward of the turbine 32 and communicates with the turbine 32. The rotationally fixed static inner support structure 16 includes a splitter 40, a static inner support housing 42 and a static outer support housing 44 located coaxial to said engine centerline A.

The axial compressor 22 includes an axial compressor rotor 46, which is mounted for rotation upon the static inner support housing 42 through an aft bearing assembly 47 and a forward bearing assembly 48. A plurality of stages of compressor blades 52 extend radially outwardly from the axial compressor rotor 46. A fixed compressor case 50 is mounted within the splitter 40. A plurality of compressor vanes 54 extend radially inwardly from the compressor case 50 between stages of the compressor blades 52. The compressor blades 52 and compressor vanes 54 are arranged circumferentially about the axial compressor rotor 46 in stages (three stages of compressor blades 52 and compressor vanes 54 are shown in this example).

A plurality of independently variable compressor inlet guide vanes 53 having pivotably mounted flaps 53A are positioned at the inlet to the axial compressor 22. Each compressor inlet guide vane includes a variable flap 53A. The flap 53A of each compressor inlet guide vane 53 is variable, i.e. it is selectively pivotable about an axis P1 that is transverse to the engine centerline. Additionally, the flap 53A of each compressor inlet guide vane 53 is pivotable independently of the flaps 53A of the other inlet guide vanes 53 or is pivotable in groups of two or more such that every flap in a group rotates together the same amount.

The rotational position of the flap 53A of each compressor inlet guide vane 53 is controlled by an independent actuator 55. The actuators 55 may be hydraulic, electric motors or any other type of suitable actuator. In the embodiment shown, the actuator 55 is located within the housing 12, radially outward of the bypass airflow path. Each actuator 55 is operatively connected to a corresponding flap 53A of an inlet guide vane via linkage, including a torque rod 56 that is routed through one of the inlet guide vanes 53. Within the splitter 40, the torque rod 56 is coupled to a trailing edge of the flap 53A via a torque rod lever 58. Within the housing 12, the actuator 55 is connected to the torque rod 56 via an actuator lever 60. Alternatively, the actuators may be directly mounted to the inner or outer end of the flap thus eliminating the linkages and torque rods.

A plurality of independently variable fan inlet guide vanes 18 having pivotably mounted flaps 18A are positioned in front of the fan blades 28. Each fan inlet guide vane 18 extends between the between the static outer support structure 14 and the static inner support structure 16 and includes a variable flap 18A. The flap 18A of each fan inlet guide vane 18 is variable, i.e. it is selectively pivotable about an axis P2 that is transverse to the engine centerline. Additionally, the flap 18A of each fan inlet guide vane 18 is pivotable independently of the flaps 18A of the other fan inlet guide vanes 18.

The rotational position of the flap 18A of each inlet guide vane is controlled by an independent actuator 115. The actuators 115 may be hydraulic, electric motors or any other type of suitable actuator. In the embodiment shown, the actuator 115 is located within the housing 12, radially outward of the bypass airflow path. Each actuator 115 is operatively connected to its corresponding flap 18A of an inlet guide vane via linkage, including a torque rod 116 that is routed through one of the fan inlet guide vanes 18. Within the splitter 40, the torque rod 116 is coupled to an outer end of the flap 18A via a torque rod lever 118. Within the housing 12, the actuator 115 is connected to the torque rod 116 via an actuator lever 120.

The fan-turbine rotor assembly 24 includes a fan hub 64 that supports a plurality of the hollow fan blades 28. Each fan blade 28 includes an inducer section 66, a hollow fan blade section 72 and a diffuser section 74. The inducer section 66 receives airflow from the axial compressor 22 generally parallel to the engine centerline A and turns the airflow from an axial airflow direction toward a radial airflow direction. The airflow is radially communicated through a core airflow passage 80 within the fan blade section 72 where the airflow is centrifugally compressed. From the core airflow passage 80, the airflow is diffused and turned once again toward an axial airflow direction toward the annular combustor 30. Preferably, the airflow is diffused axially forward in the turbine engine 10, however, the airflow may alternatively be communicated in another direction.

The tip turbine engine 10 may optionally include a gearbox assembly 90 aft of the fan-turbine rotor assembly 24, such that the fan-turbine rotor assembly 24 rotatably drives the axial compressor 22 via the gearbox assembly 90. In the embodiment shown, the gearbox assembly 90 provides a speed increase at a 3.34-to-one ratio. The gearbox assembly 90 may be an epicyclic gearbox, such as a planetary gearbox as shown, that is mounted for rotation between the static inner support housing 42 and the static outer support housing 44. The gearbox assembly 90 includes a sun gear 92, which rotates the axial compressor 22, and a planet carrier 94, which rotates with the fan-turbine rotor assembly 24. A plurality of planet gears 93 each engage the sun gear 92 and a rotationally fixed ring gear 95. The planet gears 93 are mounted to the planet carrier 94. The gearbox assembly 90 is mounted for rotation between the sun gear 92 and the static outer support housing 44 through a gearbox forward bearing 96 and a gearbox rear bearing 98. The gearbox assembly 90 may alternatively, or additionally, reverse the direction of rotation and/or may provide a decrease in rotation speed.

FIG. 2 is a schematic of three of the fan inlet guide vane flaps 18A, 18A′, 18A″ and three of the compressor inlet guide vane flaps 53A, 53A′, 53A″. The rotational position of the flap 18A, 18A′, 18A″ of each fan inlet guide vane 18, 18′, 18″ is controlled by an independent actuator 115, 115′, 115″, respectively. As is shown in FIG. 2, the torque rod 116, 116′, 116″ is connected to the flap 18A, 18A′, 18A″ via torque rod lever 118, 118′, 118″. The linkage is shown schematically in FIG. 2, but various configurations could be utilized. The actuators 115, 115′, 115″ are independently controlled by a controller or CPU 112 to selectively pivot the flaps 18A, 18A′, 18A″ to desired positions independently. For example, in FIG. 2, as controlled by the CPU 112, the first flap 18A is pivoted by actuator 115 to an angle a relative to a plane extending radially through the first flap 18A and the engine centerline A, while the second flap 18A′ is pivoted by actuator 115′ to an angle b relative to a plane through the second flap 18A′ and the engine centerline A and while the third flap 18A″ is pivoted by actuator 115″ to an angle c relative to a plane through the third flap 18A″ and the engine centerline A. Each of the angles a, b and c is varied independently of the others and can be set to different angles.

Similarly, the rotational position of the flap 53A, 53A′, 53A″ of each compressor inlet guide vane 53, 53′, 53″ is controlled by an independent actuator 55, 55′, 55″, respectively. The actuators 55, 55′, 55″ are independently controlled by CPU 112 to selectively pivot the flaps 53A, 53A′, 53A″ to desired positions independently. For example, in FIG. 2, as controlled by the CPU 112, the first flap 53A is pivoted by actuator 55 to an angle d relative to a plane through the first flap 53A and the engine centerline A, while the second flap 53A′ is pivoted by actuator 55′ to an angle e relative to a plane through the second flap 53A′ and the engine centerline A and while the third flap 53A″ is pivoted by actuator 55″ to an angle f relative to a plane through the third flap 53A″ and the engine centerline A. Each of the angles d, e and f is varied independently of the others and can be set to different angles.

In operation, referring to FIG. 1, core airflow entering the axial compressor 22 is redirected by the compressor inlet guide vanes 53 and flaps 53A before being compressed by the compressor blades 52. Selective, individual, independent variation of the compressor inlet guide vane flaps 53A control inlet distortion and increase the stability of the axial compressor 22 and the turbine engine 10. The compressed air from the axial compressor 22 enters the inducer section 66 in a direction generally parallel to the engine centerline A, and is then turned by the inducer section 66 radially outwardly through the core airflow passage 80 of the hollow fan blades 28. The airflow is further compressed centrifugally in the hollow fan blades 28 by rotation of the hollow fan blades 28. From the core airflow passage 80, the airflow is turned and diffused axially forward in the turbine engine 10 into the annular combustor 30. The compressed core airflow from the hollow fan blades 28 is mixed with fuel in the annular combustor 30 and ignited to form a high-energy gas stream.

The high-energy gas stream is expanded over the plurality of tip turbine blades 34 mounted about the outer periphery of the fan-turbine rotor assembly 24 to drive the fan-turbine rotor assembly 24, which in turn rotatably drives the axial compressor 22 either directly or via the optional gearbox assembly 90. The fan-turbine rotor assembly 24 discharges fan bypass air axially aft to merge with the core airflow from the turbine 32 in an exhaust case 106. Incoming bypass airflow is redirected by fan inlet guide vanes 18 and flaps 18A before being drawn through the fan blades 28. Selective, individual, independent variation of the fan inlet guide vane flaps 18A control inlet distortion and increase the stability of the turbine engine 10.

A plurality of exit guide vanes 108 are located between the static outer support housing 44 and the rotationally fixed static outer support structure 14 to guide the combined airflow out of the turbine engine 10 and provide forward thrust. An exhaust mixer 110 mixes the airflow from the turbine blades 34 with the bypass airflow through the fan blades 28.

FIG. 3 illustrates the turbine engine 10 of FIGS. 1-2 installed vertically in an aircraft 200. The aircraft 200 includes a conventional turbine engine 210 for primarily providing forward thrust and the turbine engine 10 for primarily providing vertical thrust. As explained above, the vertical orientation would obtain particular benefits from the individual control of the fan inlet guide vane flaps 18A and compressor inlet guide vane flaps 53A (flaps 18A and 53A are shown in FIGS. 1 and 2).

FIG. 4 illustrates an alternative variable fan inlet guide vane 218 that could be used in the turbine engine of FIGS. 1-3. The fan inlet guide vane 218 includes an interior cavity 220 leading to a plurality of fluid outlets or nozzles 222 disposed along a trailing edge and directed transversely to the surface of the fan inlet guide vane 218. Compressed air, such as bleed air from the axial compressor 22 or from the inlet to the combustor 30 (FIG. 1), is selectively supplied to each fan inlet guide vane 218, 218′, 218″ independently as controlled by an associated valve actuator 215, 215′, 215″. In this case, the linkage between the actuator 215, 215′, 215″ and the variable inlet guide vane 218 is a conduit 216, 216′, 216″. The fluid flow through the nozzles 222 redirects the incoming airflow and reduces inlet distortion, thereby improving the stability of the turbine engine 10.

Similarly, FIG. 5 illustrates an alternative variable compressor inlet guide vane 253 that could be used in the turbine engine of FIGS. 1-3. The compressor inlet guide vane 253 includes an interior cavity 254 leading to a plurality of fluid outlets or nozzles 256 aligned along a trailing edge and directed transversely to the surface of the compressor inlet guide vane 253. Compressed air, such as bleed air from the axial compressor 22 or from the inlet to the combustor 30 (FIG. 1), is selectively supplied to each compressor inlet guide vane 253, 253′, 253″ independently as controlled by an associated valve actuator 255, 255′, 255″. In this case, the linkage between the actuator 255, 255′, 255″ and the variable inlet guide vane 253, 253′, 253″ is a conduit 258, 258′, 258″. The fluid flow through the nozzles 256 redirects the incoming airflow and reduces inlet distortion, thereby improving the stability of the axial compressor 22 and the turbine engine 10.

In accordance with the provisions of the patent statutes and jurisprudence, exemplary configurations described above are considered to represent a preferred embodiment of the invention. However, it should be noted that the invention can be practiced otherwise than as specifically illustrated and described without departing from its spirit or scope. For example, there are many configurations of linkages, rigid and/or flexible, that could be used to connect the actuator 115 to the inlet guide vane flaps 18A. Also, although the actuator 115 has been shown in connection with a tip turbine engine 10, it could also be used in conventional or other turbine engines. Although the invention has been shown with a single actuator 115 for each inlet guide vane flap 18A, it is also possible that one actuator 115 could control more than one inlet guide vane flap 18A.

Claims

1. A method for controlling a plurality of inlet guide vanes of a turbine engine, the method including the steps of:

a) varying a first inlet guide vane of the plurality of inlet guide vanes to a first amount with a first actuator; and
b) varying a second inlet guide vane of the plurality of inlet guide vanes to a second amount with a second actuator while the first inlet guide vane is at the first amount, the first amount being different from the second amount, wherein the first actuator and the second actuator each independently control only one inlet guide vane.

2. The method of claim 1 wherein said step a) further includes the step of pivoting the first inlet guide vane to a first angle relative to a longitudinal axis through the turbine engine, and said step b) further includes pivoting the second inlet guide vane to a second angle relative to the longitudinal axis while the first inlet guide vane is at the first angle, the first angle being different from the second angle.

3. The method of claim 2 further including the step of varying the first angle and the second angle independently of one another.

4. The method of claim 1 wherein the plurality of inlet guide vanes are located radially inward of a bypass air flow path.

5. The method of claim 1 wherein the plurality of inlet guide vanes are mounted in a bypass air flow path.

6. The method of claim 1 wherein the first inlet guide vane and the second inlet guide vane each include at least one fluid outlet, said step a) including the step of varying fluid flow through the at least one fluid outlet in the first inlet guide vane, said step b) including the step of varying fluid flow through the at least one fluid outlet in the second inlet guide vane.

7. The method of claim 1, wherein the first actuator controls only the first inlet guide vane.

8. The method of claim 1, wherein the second actuator controls only the second inlet guide vane.

Referenced Cited
U.S. Patent Documents
1544318 June 1925 Hodgkinson
2221685 November 1940 Smith
2414410 January 1947 Griffith
2499831 March 1950 Palmatier
2548975 April 1951 Hawthorne
2611241 September 1952 Schulz
2620554 December 1952 Mochel et al.
2698711 January 1955 Newcomb
2801789 August 1957 Moss
2830754 April 1958 Stalker
2874926 February 1959 Gaubatz
2989848 June 1961 Paiement
3009630 November 1961 Busquet
3037742 June 1962 Dent et al.
3042349 July 1962 Pirtle et al.
3081597 March 1963 Kosin et al.
3132842 May 1964 Tharp
3204401 September 1965 Serriades
3216455 November 1965 Cornell et al.
3267667 August 1966 Erwin
3269120 August 1966 Sabatiuk
3283509 November 1966 Nitsch
3286461 November 1966 Johnson
3302397 February 1967 Davidovic
3363419 January 1968 Wilde
3404831 October 1968 Campbell
3465526 September 1969 Emerick
3496725 February 1970 Ferri et al.
3505819 April 1970 Wilde
3616616 November 1971 Flatt
3684857 August 1972 Morley et al.
3703081 November 1972 Krebs et al.
3705775 December 1972 Rioux
3720060 March 1973 Davies et al.
3729957 May 1973 Petrie et al.
3735593 May 1973 Howell
3811273 May 1974 Martin
3818695 June 1974 Rylewski
3836279 September 1974 Lee
3861822 January 1975 Wanger
3932813 January 13, 1976 Gallant
3979087 September 7, 1976 Boris et al.
4005575 February 1, 1977 Scott et al.
4043121 August 23, 1977 Thomas et al.
4130379 December 19, 1978 Partington
4147035 April 3, 1979 Moore et al.
4193738 March 18, 1980 Landis, Jr. et al.
4251185 February 17, 1981 Karstensen
4251987 February 24, 1981 Adamson
4265646 May 5, 1981 Weinstein et al.
4271674 June 9, 1981 Marshall et al.
4298090 November 3, 1981 Chapman
4314791 February 9, 1982 Weiler
4326682 April 27, 1982 Nightingale
4452038 June 5, 1984 Soligny
4463553 August 7, 1984 Boudigues
4561257 December 31, 1985 Kwan et al.
4563875 January 14, 1986 Howald
4631092 December 23, 1986 Ruckle et al.
4751816 June 21, 1988 Perry
4785625 November 22, 1988 Stryker et al.
4817382 April 4, 1989 Rudolph et al.
4834614 May 30, 1989 Davids et al.
4883404 November 28, 1989 Sherman
4887424 December 19, 1989 Geidel et al.
4904160 February 27, 1990 Partington
4912927 April 3, 1990 Billington
4965994 October 30, 1990 Ciokajlo et al.
4999994 March 19, 1991 Rud et al.
5010729 April 30, 1991 Adamson et al.
5012640 May 7, 1991 Mirville
5014508 May 14, 1991 Lifka
5088742 February 18, 1992 Catlow
5107676 April 28, 1992 Hadaway et al.
5157915 October 27, 1992 Bart
5182906 February 2, 1993 Gilchrist et al.
5224339 July 6, 1993 Hayes
5232333 August 3, 1993 Girault
5267397 December 7, 1993 Wilcox
5269139 December 14, 1993 Klees
5275536 January 4, 1994 Stephens et al.
5315821 May 31, 1994 Dunbar et al.
5328324 July 12, 1994 Dodd
5443590 August 22, 1995 Ciokajlo et al.
5466198 November 14, 1995 McKibbin et al.
5472314 December 5, 1995 Delonge et al.
5497961 March 12, 1996 Newton
5501575 March 26, 1996 Eldredge et al.
5537814 July 23, 1996 Nastuk et al.
5584660 December 17, 1996 Carter et al.
5628621 May 13, 1997 Toborg
5746391 May 5, 1998 Rodgers et al.
5769317 June 23, 1998 Sokhey et al.
6004095 December 21, 1999 Waitz et al.
6095750 August 1, 2000 Ross et al.
6102361 August 15, 2000 Riikonen
6158207 December 12, 2000 Polenick et al.
6223616 May 1, 2001 Sheridan
6244539 June 12, 2001 Lifson et al.
6254346 July 3, 2001 Fukuno et al.
6364805 April 2, 2002 Stegherr
6381948 May 7, 2002 Klingels
6382915 May 7, 2002 Aschermann et al.
6384494 May 7, 2002 Avidano et al.
6430917 August 13, 2002 Platts
6454535 September 24, 2002 Goshorn et al.
6471474 October 29, 2002 Mielke et al.
RE37900 November 5, 2002 Partington
6513334 February 4, 2003 Varney
6619030 September 16, 2003 Seda et al.
6851264 February 8, 2005 Kirtley et al.
6883303 April 26, 2005 Seda
6910854 June 28, 2005 Joslin
7021042 April 4, 2006 Law
7214157 May 8, 2007 Flamang et al.
7854112 December 21, 2010 Roberge
7934902 May 3, 2011 Suciu et al.
7976272 July 12, 2011 Suciu et al.
7980054 July 19, 2011 Suciu et al.
8276362 October 2, 2012 Suciu et al.
8641367 February 4, 2014 Norris et al.
20020190139 December 19, 2002 Morrison
20030031556 February 13, 2003 Mulcaire et al.
20030131602 July 17, 2003 Ingistov
20030131607 July 17, 2003 Daggett
20030161724 August 28, 2003 Capozzi et al.
20030192303 October 16, 2003 Paul
20030192304 October 16, 2003 Paul
20040025490 February 12, 2004 Paul
20040070211 April 15, 2004 Franchet et al.
20040189108 September 30, 2004 Dooley
20040219024 November 4, 2004 Soupizon et al.
20050008476 January 13, 2005 Eleftheriou
20050127905 June 16, 2005 Proctor et al.
20090145105 June 11, 2009 Suciu et al.
20090148273 June 11, 2009 Suciu et al.
20090232643 September 17, 2009 Norris et al.
Foreign Patent Documents
2451059 April 1976 DE
19646601 April 1997 DE
2599086 November 1987 FR
907323 October 1962 GB
1287223 August 1972 GB
1351000 April 1974 GB
1357016 June 1974 GB
1466613 March 1977 GB
2016597 September 1979 GB
2026102 January 1980 GB
2095755 October 1982 GB
2191606 December 1987 GB
2265221 September 1993 GB
2410530 August 2005 GB
10184305 July 1998 JP
02081883 October 2002 WO
2004092567 October 2004 WO
Other references
  • PCT International Search Report for PCT/US2004/039974 date of completion Feb. 16, 2006.
Patent History
Patent number: 8967945
Type: Grant
Filed: Dec 23, 2013
Date of Patent: Mar 3, 2015
Patent Publication Number: 20140219772
Assignee: United Technologies Corporation (Hartford, CT)
Inventors: Craig A. Nordeen (Manchester, CT), James W. Norris (Lebanon, CT)
Primary Examiner: Christopher Verdier
Application Number: 14/138,889